Aircraft are assembled from a number of separate assemblies which are fixedly mounted to each other during assembly of the aircraft. Such an assembly includes wing boxes, leading and trailing edges of a wing, horizontal tail planes (HTPs), Vertical Tail Planes (VTPs) and the fuselage, which are themselves formed from a number of discrete components including spars, ribs, skin, gunnels, compression struts, and stringers which are positioned relative to and fixedly mounted to each other.
Such components and subassemblies have stringent tolerances and are generally fixedly located relative to each other using conventional mechanical fixings, such as bolts and rivets, received in pre-formed holes extending through a section of each component to be fixedly located. Although such conventional fixings form a secure joint to fixedly locate adjacent surfaces, they are known to have a number of disadvantages when used in an aerospace assembly.
Conventional fixings are known to cause metal fatigue and localised stress concentrations of the area proximal to each pre-formed hole, which may lead to a structural failure of the joint and/or the components, and cause, for example, fuel leakage or overall breakdown of the aircraft.
Furthermore, mechanical fixtures made of metal may in particular pose a problem to aircrafts formed from carbon fibre composites if the aircraft was to be struck by lightening, as this may cause electrical failure or ignition of the fuel held in the fuel tank.
There has been a move within the aerospace industry towards components formed of carbon fibre materials, which have certain weight and strength advantages over their to metallic counterparts, such as aluminium sheet, moulding or extrusion. One issue with components formed from carbon fibre composite materials is that although one surface is generally produced to an exact tolerance, the opposing surface is generally of low tolerance, typically up to +/−4%, depending on the process used. With conventional fixings it is not possible for these inevitable manufacturing tolerances to be taken up is during assembly, and so it is necessary to modify or shim components to obtain the exact tolerances needed to ensure the desired joint, which is a complex and time-consuming process.
Irrespective of the manufacturing technique of aircrafts, there will be gaps and voids between the structural components as they are assembled. These gaps may be filled with shims or post machining adjustment techniques may be employed to meet the close tolerances of the components. Consequently, the manufacturing of aircraft is time consuming and usually results in wastage and a high scrap rate such that it is difficult to manufacture and assembly high volumes of aircrafts and aeronautical components, whilst also meeting the requirements for close tolerance, which is particularly difficult when utilising carbon fibre composites. It is therefore desirable to provide an apparatus which allows for the inevitable manufacturing misalignments generated from manufacturing tolerances to be taken up.
The present invention therefore seeks to provide a device for locating a first aerospace component relative to a second aerospace component which substantially overcomes or alleviates the known problems discussed above.